Engineering topics

Ghaffari, Farhad

Publications and source records attributed to Ghaffari, Farhad.

10 records

NASA NTRS2012

Error Estimates of the Ares I Computed Turbulent Ascent Longitudinal Aerodynamic Analysis

Numerical predictions of the longitudinal aerodynamic characteristics for the Ares I class of vehicles, along with the associated error estimate derived from an iterative convergence grid refinement, are presented. Computational results are based on an unstructured grid, Reynolds-averaged Navier-Stokes analysis. The validity of the approach to compute the associated error estimates, derived from a base grid to an extrapolated infinite-size grid, was first demonstrated on a sub-scaled wind tunnel model at representative ascent flow conditions for which the experimental data existed. Such analysis at the transonic flow conditions revealed a maximum deviation of about 23% between the computed longitudinal aerodynamic coefficients with the base grid and the measured data across the entire roll angles. This maximum deviation from the wind tunnel data was associated with the computed normal force coefficient at the transonic flow condition and was reduced to approximately 16% based on the infinite-size grid. However, all the computed aerodynamic coefficients with the base grid at the supersonic flow conditions showed a maximum deviation of only about 8% with that level being improved to approximately 5% for the infinite-size grid. The results and the error estimates based on the established procedure are also presented for the flight flow conditions.

Abdol-Hamid, Khaled S.↗

NASA NTRS2005

Turbulent Vortex-Flow Simulation Over a 65 deg Sharp and Blunt Leading-Edge Delta Wing at Subsonic Speeds

Turbulent thin-layer, Reynolds-Averaged Navier-Stokes solutions, based on a multi-block structured grid, are presented for a 65 deg delta wing having either a sharp leading edge (SLE) or blunt leading edge (BLE) geometry. The primary objective of the study is to assess the prediction capability of the method for simulating the leading-edge flow separation and the ensuing vortex flow characteristics. Computational results are obtained for two angles of attack of approximately 13 and 20 deg, at free-stream Mach number of 0.40 and Reynolds number of 6 million based on the wing mean aerodynamic chord. The effects of two turbulence models of Baldwin-Lomax with Degani-Schiff (BL/DS) and the Spalart-Allmaras (SA) on the numerical results are also discussed. The computations also explore the effects of two numerical flux-splitting schemes, i.e., flux difference splitting (fds) and flux vector splitting (fvs), on the solution development and convergence characteristics. The resulting trends in solution sensitivity to grid resolution for the selected leading-edge geometries, angles of attack, turbulence models and flux splitting schemes are also presented. The validity of the numerical results is evaluated against a unique set of experimental wind-tunnel data that was obtained in the National Transonic Facility at the NASA Langley Research Center.

Ghaffari, Farhad↗

NASA NTRS2002

Turbulent Navier-Stokes Flow Analysis of an Advanced Semispan Diamond-Wing Model in Tunnel and Free Air at High-Lift Conditions

Turbulent Navier-Stokes computational results are presented for an advanced diamond wing semispan model at low-speed, high-lift conditions. The numerical results are obtained in support of a wind-tunnel test that was conducted in the National Transonic Facility at the NASA Langley Research Center. The model incorporated a generic fuselage and was mounted on the tunnel sidewall using a constant-width non-metric standoff. The computations were performed at to a nominal approach and landing flow conditions.The computed high-lift flow characteristics for the model in both the tunnel and in free-air environment are presented. The computed wing pressure distributions agreed well with the measured data and they both indicated a small effect due to the tunnel wall interference effects. However, the wall interference effects were found to be relatively more pronounced in the measured and the computed lift, drag and pitching moment. Although the magnitudes of the computed forces and moment were slightly off compared to the measured data, the increments due the wall interference effects were predicted reasonably well. The numerical results are also presented on the combined effects of the tunnel sidewall boundary layer and the standoff geometry on the fuselage forebody pressure distributions and the resulting impact on the configuration longitudinal aerodynamic characteristics.

Ghaffari, Farhad↗

NASA NTRS1999

Unstructured Grid Euler Method Assessment for Longitudinal and Lateral/Directional Stability Analysis of the HSR Reference H Configuration at Transonic Speeds

Transonic Euler computations, based on unstructured grid methodology, are performed for a proposed High Speed Civil Transport (HSCT) configuration, designated as the Reference H configuration within the High Speed Research (HSR) Program. The predicted results are correlated with appropriate experimental wind-tunnel data for the baseline configuration with and without control surface deflections for a range of angle of attack at M(sub infinity) = 0.95. Good correlations between the predictions and measured data have been obtained for the longitudinal aerodynamic characteristics of the baseline configuration. The incremental effects in the longitudinal aerodynamic characteristics due to horizontal rail deflections as well as wing leading-edge and trailing-edge flap deflections have also been predicted reasonably well. Computational results and correlations with data are also presented for the lateral and directional stability characteristics for a range of angle of attack at a constant sideslip angle as well as a range of sideslip angles at a constant angle of attack. In addition, the results are presented to assess the computational method performance and convergence characteristics.

Ghaffari, Farhad↗

NASA NTRS1999

Unstructured Grid Euler Method Assessment for Longitudinal and Lateral/Directional Aerodynamic Performance Analysis of the HSR Technology Concept Airplane at Supersonic Cruise Speed

Unstructured grid Euler computations, performed at supersonic cruise speed, are presented for a High Speed Civil Transport (HSCT) configuration, designated as the Technology Concept Airplane (TCA) within the High Speed Research (HSR) Program. The numerical results are obtained for the complete TCA cruise configuration which includes the wing, fuselage, empennage, diverters, and flow through nacelles at M (sub infinity) = 2.4 for a range of angles-of-attack and sideslip. Although all the present computations are performed for the complete TCA configuration, appropriate assumptions derived from the fundamental supersonic aerodynamic principles have been made to extract aerodynamic predictions to complement the experimental data obtained from a 1.675%-scaled truncated (aft fuselage/empennage components removed) TCA model. The validity of the computational results, derived from the latter assumptions, are thoroughly addressed and discussed in detail. The computed surface and off-surface flow characteristics are analyzed and the pressure coefficient contours on the wing lower surface are shown to correlate reasonably well with the available pressure sensitive paint results, particularly, for the complex flow structures around the nacelles. The predicted longitudinal and lateral/directional performance characteristics for the truncated TCA configuration are shown to correlate very well with the corresponding wind-tunnel data across the examined range of angles-of-attack and sideslip. The complementary computational results for the longitudinal and lateral/directional performance characteristics for the complete TCA configuration are also presented along with the aerodynamic effects due to empennage components. Results are also presented to assess the computational method performance, solution sensitivity to grid refinement, and solution convergence characteristics.

Ghaffari, Farhad↗

NASA NTRS1999

Unstructured Grid Viscous Flow Simulation Over High-Speed Research Technology Concept Airplane at High-Lift Conditions

Numerical viscous solutions based on an unstructured grid methodology are presented for a candidate high-speed civil transport configuration, designated as the Technology Concept Airplane (TCA), within the High-Speed Research (HSR) program. The numerical results are obtained on a representative TCA high-lift configuration that consisted of the fuselage and the wing, with deflected full-span leading-edge and trailing-edge flaps. Typical on-and off-surface flow structures, computed at high-lift conditions appropriate for the takeoff and landing, indicated features that are generally plausible. Reasonable surface pressure correlations between the numerical results and the experimental data are obtained at free-stream Mach number M(sub infinity) = 0.25 and Reynolds number based on bar-c R(sub c) = 8 x 10(exp 6) for moderate angles of attack of 9.7 deg. and 13.5 deg. However, above and below this angle-of-attack range, the correlation between computed and measured pressure distributions starts to deteriorate over the examined angle-of-attack range. The predicted longitudinal aerodynamic characteristics are shown to correlate very well with existing experimental data across the examined angle-of-attack range. An excellent agreement is also obtained between the predicted lift-to-drag ratio and the experimental data over the examined range of flow conditions.

Ghaffari, Farhad↗

NASA NTRS1994

Navier-Stokes, flight, and wind tunnel flow analysis for the F/A-18 aircraft

Computational analysis of flow over the F/A-18 aircraft is presented along with complementary data from both flight and wind tunnel experiments. The computational results are based on the three-dimensional thin-layer Navier-Stokes formulation and are obtained from an accurate surface representation of the fuselage, leading-edge extension (LEX), and the wing geometry. However, the constraints imposed by either the flow solver and/or the complexity associated with the flow-field grid generation required certain geometrical approximations to be implemented in the present numerical model. In particular, such constraints inspired the removal of the empennage and the blocking (fairing) of the inlet face. The results are computed for three different free-stream flow conditions and compared with flight test data of surface pressure coefficients, surface tuft flow, and off-surface vortical flow characteristics that included breakdown phenomena. Excellent surface pressure coefficient correlations, both in terms of magnitude and overall trend, are obtained on the forebody throughout the range of flow conditions. Reasonable pressure agreement was obtained over the LEX; the general correlation tends to improve at higher angles of attack. The surface tuft flow and the off-surface vortex flow structures compared qualitatively well with the flight test results. To evaluate the computational results, a wind tunnel investigation was conducted to determine the effects of existing configurational differences between the flight vehicle and the numerical model on aerodynamic characteristics. In most cases, the geometrical approximations made to the numerical model had very little effect on overall aerodynamic characteristics.

Ghaffari, Farhad↗

NASA NTRS1991

Navier-Stokes solutions about the F/A-18 wing-LEX-fuselage configuration with multi-block structured grids

Three-dimensional thin-layer Navier-Stokes computations are presented for the F/A-18 configuration. The modeled configuration includes an accurate surface representation of the fuselage, leading-edge-extension, as well as the wing with and without leading-edge-flap deflection. A multi-block structured volume grid with various topologies is generated using transfinite interpolation technique. The flowfield domain is divided into twenty blocks, each representing a particular geometrical complexity of the configuration. The results are obtained from an algorithm for solving the compressible Navier-Stokes equations that incorporates an upwind-biased, flux-difference-splitting approach. In addition, a newly developed capability that allows for generalized surface patching among blocks is employed. Turbulent results are presented for flow conditions that correspond to recent NASA F/A-18 High Alpha Research Vehicle flight experiments. Good correlations between the computations and the flight test results are disclosed for both surface flow patterns as well as surface pressure distributions.

Ghaffari, Farhad↗

NASA NTRS1990

Transonic Navier-Stokes solutions about a complex high-speed accelerator configuration

Three-dimensional transonic viscous flow computations are presented for a generic high-speed accelerator model which includes wing, body, fillets, and a no-flow through engine nacelle. Solutions are obtained from an algorithm for the compressible Navier-Stokes equations which incorporates an upwind-biased, flux-vector-splitting approach along with longitudinally-patched grids. Results are presented for fully turbulent flow assumptions and include correlations with wind tunnel data. A good quantitative agreement for the forebody surface pressure distribution is achieved between computations and the available wind-tunnel measurements at freestream Mach number equal to 0.9. Furthermore, it is demonstrated that the flow is stagnating around the boattail region due to separation from the aft-engine cowl lip.

Ghaffari, Farhad↗

NASA NTRS1989

A patched-grid algorithm for complex configurations directed towards the F-18 aircraft

A patched-grid algorithm for the analysis of complex configurations with an implicit, upwind-biased Navier-Stokes solver is presented. Results from both a spatial-flux and a time-flux conservation approach to patching across zonal boundaries are presented. A generalized coordinate transformation with a biquadratic geometric element is used at the zonal interface in order to treat highly stretched viscous grids and arbitrarily-shaped zonal boundaries. Applications are made to the F-18 forebody-strake configuration at subsonic, high-alpha conditions. Computed surface flow patterns compare well with ground-based and flight-test results; the large effect of Reynolds number on the forebody flowfield is shown.

Thomas, James L.↗